U.S. patent number 8,646,279 [Application Number 13/480,696] was granted by the patent office on 2014-02-11 for segment component in high-temperature casting material for an annular combustion chamber, annular combustion chamber for an aircraft engine, aircraft engine and method for the manufacture of an annular combustion chamber.
This patent grant is currently assigned to Rolls-Royce Deutschland Ltd & Co KG. The grantee listed for this patent is Miklos Gerendas, Karl Schreiber. Invention is credited to Miklos Gerendas, Karl Schreiber.
United States Patent |
8,646,279 |
Schreiber , et al. |
February 11, 2014 |
Segment component in high-temperature casting material for an
annular combustion chamber, annular combustion chamber for an
aircraft engine, aircraft engine and method for the manufacture of
an annular combustion chamber
Abstract
The present invention relates to a segment component in
high-temperature casting material for an annular combustion chamber
of an aircraft engine, characterized by a combustion-chamber wall
which in operation shields a fuel flame extending along a burner
axis from the environment, with the combustion-chamber wall having
a bulge which points in a direction facing away from the burner
axis. The invention furthermore relates to an annular combustion
chamber, an aircraft engine with an annular combustion chamber as
well as a method for the manufacture of an annular combustion
chamber.
Inventors: |
Schreiber; Karl (Am Mellensee,
DE), Gerendas; Miklos (Am Mellensee, DE) |
Applicant: |
Name |
City |
State |
Country |
Type |
Schreiber; Karl
Gerendas; Miklos |
Am Mellensee
Am Mellensee |
N/A
N/A |
DE
DE |
|
|
Assignee: |
Rolls-Royce Deutschland Ltd &
Co KG (DE)
|
Family
ID: |
46148713 |
Appl.
No.: |
13/480,696 |
Filed: |
May 25, 2012 |
Prior Publication Data
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Document
Identifier |
Publication Date |
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US 20120304658 A1 |
Dec 6, 2012 |
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Foreign Application Priority Data
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May 25, 2011 [DE] |
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10 2011 076 473 |
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Current U.S.
Class: |
60/804;
60/752 |
Current CPC
Class: |
F23R
3/50 (20130101); Y10T 29/4927 (20150115); F23R
2900/00018 (20130101) |
Current International
Class: |
F02C
3/06 (20060101); F02C 1/00 (20060101); F02G
3/00 (20060101) |
Field of
Search: |
;60/752,754-758,760,804 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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19959292 |
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Jun 2001 |
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DE |
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1106927 |
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Jun 2001 |
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EP |
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1312865 |
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May 2003 |
|
EP |
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1422479 |
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May 2004 |
|
EP |
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2004/106809 |
|
Dec 2004 |
|
WO |
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2009/109409 |
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Sep 2009 |
|
WO |
|
Other References
Nickel based superalloy welding practices for industrial gas
turbine applications Authors: Henderson, M. B.; Arrell, D.;
Larsson, R.; Heobel, M.; Marchant, G. Source: Science and
Technology of Welding & Joining, vol. 9, No. 1, Feb. 2004 , pp.
13-21(9). cited by examiner.
|
Primary Examiner: Wongwian; Phutthiwat
Assistant Examiner: Breazeal; William
Attorney, Agent or Firm: Klima; Timothy J. Shuttleworth
& Ingersoll, PLC
Claims
What is claimed is:
1. A segment component in high-temperature casting material for an
annular combustion chamber of an aircraft engine, the segment
component forming a combustion-chamber wall for shielding a fuel
flame extending along a burner axis from the environment; the
combustion-chamber wall having a bulge bulging in a direction away
from the burner axis; the segment component including a
combustion-chamber head, and an inner combustion-chamber wall and
an outer combustion-chamber wall between which the fuel flame is
provided in operation, with the inner combustion-chamber wall, the
combustion-chamber head and the outer combustion-chamber wall being
connected to one another as a monolithic, U-shaped casting.
2. The segment component of claim 1, wherein at least one chosen
from the inner combustion-chamber wall and the outer
combustion-chamber wall includes the bulge.
3. The segment component of claim 1, wherein the bulge of the
combustion-chamber wall is adapted substantially to a contour of
the fuel flame in operation, with at least one chosen from a length
of the bulge (L.sub.B) and a width of the bulge (B.sub.B)
corresponding substantially to a respective length and width of the
fuel flame in operation.
4. The segment component of claim 1, wherein the high-temperature
casting material is a super-alloy containing at least one chosen
from nickel, chromium, cobalt, nickel-iron, Inconel 738, Inconel
738 LC, Inconel 939, Inconel 939 LC, Inconel 713, Inconel 713 LC,
C1023, Mar M 002, and CM 274LC.
5. The segment component of claim 4, and further comprising at
least one chosen from a mounting flange and a device for arranging
an injector is provided on the combustion-chamber head.
6. The segment component of claim 1, and further comprising at
least one nozzle for cooling air integrally formed onto a
combustion-chamber wall.
7. The segment component of claim 1, wherein the combustion-chamber
wall has a mean thickness between 1 and 4 mm.
8. The segment component of claim 7, wherein the combustion-chamber
wall has a mean thickness between 1.4 to 3 mm.
9. An annular combustion chamber for an aircraft engine,
comprising: at least two segment components, each segment component
in high-temperature casting material for an annular combustion
chamber of an aircraft engine, the segment component forming a
combustion chamber wall which in operation shields a fuel flame
extending along a burner axis from the environment; the
combustion-chamber wall having a bulge bulging in a direction
facing away from the burner axis; each segment component including
a combustion-chamber head, and an inner combustion-chamber wall and
an outer combustion-chamber wall between which the fuel flame is
provided in operation, with the inner combustion-chamber wall, the
combustion-chamber head and the outer combustion-chamber wall being
connected to one another as a monolithic, U-shaped casting.
10. The annular combustion chamber in accordance with claim 9, and
further comprising a varying annular space height along a
circumference of an annular space between the inner
combustion-chamber wall and the outer combustion-chamber wall with
areas having a greater annular space height alternating with areas
having a lower annular space height along the circumference.
11. The annular combustion chamber in accordance with claim 9, and
further comprising a varying annular space height along a
circumference of an annular space between the inner
combustion-chamber wall and the outer combustion-chamber wall with
areas having a greater annular space height and areas having a
lower annular space height, with injectors for the fuel being
provided in the areas with the greater annular space height.
12. An aircraft engine including an annular combustion chamber in
accordance with claim 9.
13. A method for manufacturing an annular combustion chamber for an
aircraft engine, comprising: casting each of at least two segment
components from a high-temperature casting material, each of the at
least two segment components including an inner combustion-chamber
wall, an outer combustion-chamber wall and a combustion-chamber
head portion connected to one another as a monolithic, U-shaped
casting, with at least one chosen from the inner combustion-chamber
wall and the outer combustion-chamber cast to include a bulge
bulging in a direction facing away from a burner axis; subsequently
welding the at least two segment components together to form the
annular combustion chamber.
14. The method of claim 13, wherein the segment components are
connected to one another by at least one chosen from electron beam
welding and laser welding.
15. The method of claim 14, wherein the welding is performed with
at least one chosen from IN626 filler, Polymet 972 and other
ductile filler materials.
Description
This application claims priority to German Patent Application
DE102011076473.9 filed May 25, 2011, the entirety of which is
incorporated by reference herein.
This invention relates to a segment component in high-temperature
casting material for an annular combustion chamber, an annular
combustion chamber for an aircraft engine, an aircraft engine and a
method for the manufacture of an annular combustion chamber.
Modern aircraft engines usually have annular combustion chambers
arranged axially between the compressor and the turbine. An annular
combustion chamber has, coaxially to the engine longitudinal axis,
an annular space delimited by combustion-chamber walls and referred
to as flame tube. The injectors for the fuel are arranged along the
annular cross-section of the annular space. In operation, the fuel
flames extend from these injectors into the annular space.
Due to the high thermal loads, the combustion-chamber walls must be
designed with adequate thermal stability. It is known for example,
to equip the combustion-chamber walls with particularly
thermo-resistant plates. A method is known from EP 1 106 927
according to which the annular space of an annular combustion
chamber is made up of individual segments of casting material, with
high-temperature casting materials being used.
The object underlying the present invention is to provide segment
components for annular combustion chambers which are thermically
and fluidically improved.
It is an object of the present invention to provide a solution to
the above problems by a segment component having features as
described herein.
In this case, a combustion-chamber wall which in operation shields
a fuel flame extending along a burner axis from the environment has
a bulge which points in a direction facing away from the burner
axis. A part of a segment component for an outer combustion-chamber
wall of an annular combustion chamber has for example a bulge
pointing radially outwards. A part of a segment component for an
inner combustion-chamber wall has for example a bulge which points
outwards. The bulges create in the immediate vicinity of the burner
flame a larger space, in that the spacing of the combustion-chamber
walls is increased in at least some areas around the burner
flame.
It is advantageous here to use an inner combustion-chamber wall and
an outer combustion-chamber wall, between which a fuel flame is
provided along a burner axis in operation, and which for example
feature a U-shaped arrangement. The inner and/or the outer
combustion-chamber wall then have a bulge in the direction pointing
away from the burner axis.
It is particularly advantageous here when the at least one bulge of
the combustion-chamber wall is adapted substantially to the contour
of the fuel flame in operation. The length and/or width of the
bulge can here advantageously correspond substantially to the
length and/or width of the fuel flame in operation.
Advantageous high-temperature casting materials are a super-alloy
containing nickel, chromium, cobalt and/or nickel-iron, in
particular Inconel 738/Inconel 738 LC, Inconel 939/Inconel 939 LC,
Inconel 713/Inconel 713 LC, C1023, Mar M 002 and/or CM 274LC. These
materials have a sufficient temperature resistance.
In an advantageous embodiment, the inner combustion-chamber wall
and the outer combustion-chamber wall are connected to one another
in one piece as a casting by a combustion-chamber head, or the
inner combustion-chamber wall and the outer combustion-chamber wall
are connected to a combustion-chamber head. In the first variant,
one-piece segment components are provided, and in the second
variant two segment components connected to one another are
provided.
An advantageous embodiment is obtained when at least one mounting
flange is arranged on the combustion-chamber head. It is
furthermore advantageous when a device for arranging an injector
for fuel is provided on the combustion-chamber head. At least one
nozzle for cooling air integrally formed onto a combustion-chamber
wall can also be advantageously provided.
The combustion-chamber wall advantageously has in one embodiment a
mean thickness between 1 and 4 mm, in particular 1.4 to 3 mm.
The problem is resolved by providing an annular combustion chamber
for an aircraft engine having the features of Claim 9. For this
purpose at least two segment components in accordance with at least
one of the Claims 1 to 8 are used.
Advantageous embodiments of the annular combustion chamber have a
variable annular space height along the circumference of the
annular space. By adapting the annular space height to, for
example, burner flames and/or injectors, the thermal and/or
mechanical load of the walls can be attained. This applies in
particular when areas A with a greater annular space height
H.sub.RA alternate with areas B with a lower annular space height
H.sub.RB along the circumference, such that the combustion-chamber
walls form a kind of wavelike structure.
It is particularly advantageous here when areas with a greater
annular space height and areas with a lower annular space height
are formed, where during assembly injectors for the fuel are
provided in the areas with the greater annular space height. The
areas with greater annular space height give the fuel flame more
space and shield it from disturbances inside the annular space.
Furthermore, the segment components are in advantageous embodiments
connected to one another by welds, in particular electron beam
welds, laser welds with IN626 Filler, Polymet 972 or other ductile
filler materials.
The problem is also resolved by providing an aircraft engine with
an annular combustion chamber in accordance with the Claims 11 to
14. The entire flow from the compressor via the combustion chamber
to the turbine is improved by the bulges arranged around the
flames.
Furthermore, the problem is resolved by a method for the
manufacture of an annular combustion chamber.
In one embodiment, at least two segment components are cast with an
inner combustion-chamber wall, an outer combustion-chamber wall and
a combustion-chamber head from high-temperature casting material.
The at least two segment components are subsequently connected by
joining them, in particular by welding, to the annular combustion
chamber.
Alternatively, at least two segment components are connected, in
particular welded, to form an inner full ring structure. At least
two segment components are connected, in particular welded, to form
an outer full ring structure. The present full ring structures are
connected to a combustion-chamber head structure.
The invention is described in greater detail in the following with
reference to the figures of the accompanying drawing showing
several exemplary embodiments. In the drawing,
FIG. 1 shows a schematic perspective representation of an annular
combustion chamber known per se,
FIG. 2 shows a perspective representation of an embodiment of a
segment component with two combustion-chamber walls for an annular
combustion chamber,
FIG. 2A shows a view from the combustion-chamber head onto the
embodiment as per FIG. 2,
FIG. 2B shows a sectional view of the embodiment as per FIG. 2 in
the longitudinal direction,
FIG. 2C shows a sectional view of the embodiment as per FIG. 2,
perpendicularly to the longitudinal direction,
FIG. 3 shows an axial sectional view onto an embodiment for an
annular combustion chamber formed by segment components in
accordance with the embodiment as per FIG. 2,
FIG. 4 shows a top view onto a further embodiment of a segment
component with two combustion-chamber walls,
FIG. 5 shows a further embodiment of a segment component with a
combustion-chamber wall,
FIG. 6A shows a perspective view of a first stage of an annular
space structure,
FIG. 6B shows a perspective view of a second stage of an annular
space structure.
FIG. 1 shows in a perspective view an annular combustion chamber
with an annular space 30, as used for example in an aircraft
engine.
The annular space 30 is arranged in the main flow direction of the
aircraft engine downstream of the compressor (not shown here) and
the intake area of a turbine 40. In the representation of FIG. 1,
two injectors 25 are visible, from which fuel flames 20 (not shown
here) emanate along burner axes 21 during operation. The burner
axes 21 and hence also the fuel flames 20 are thus between the
inner combustion-chamber wall 11 and the outer combustion-chamber
wall 12. This annular space 30 is also referred to as flame tube.
The combustion-chamber walls 11, 12 thus shield the fuel flames 20
inwardly and outwardly from the environment.
The distance between the combustion-chamber walls 11, 12, the
annular space height H.sub.R (also referred to as flame space
height), varies in the axial direction of the aircraft engine, but
is constant along the circumference of the annular combustion
chamber 10.
The invention described in the following on the basis of various
embodiments relates among others to annular combustion chambers
where the annular combustion chamber height H.sub.R is non-constant
along the circumference.
An annular combustion chamber of this type is for example made up
of at least two segment components 10 of high-temperature casting
material. In the case of two segment components, each of the
segment components 10 provides for example 180.degree. of the
annular space 30.
FIG. 2 shows a segment component 10 covering a considerably smaller
angular area, i.e. 30.degree., as can be discerned particularly
clearly from the view of FIG. 2A.
An annular combustion chamber composed of such segment components
10 thus has twelve of these segment components 10. In principle it
is possible to design the segment components 10 with a different
geometry, so that fewer or more than twelve segment components 10
are used. Here too it is not essential that an even number of
segment components 10 is used to form an annular space 30.
FIG. 2 shows an embodiment of a segment component 10 in which parts
form the inner combustion-chamber wall 11 and the outer
combustion-chamber wall 12 when the segment components 10 are put
together (see FIG. 5). An opening 24 for the injector 25 (not shown
here) is provided on the combustion-chamber head 22. The fuel flame
20 (not shown here) created with the injector 25 extends along the
burner axis 21 into the annular space 30 and in the direction of
the intake area of the turbine 40 (not shown here, see FIG. 1).
This embodiment of the segment component 10 is made in one piece
from a high-temperature casting material. A super-alloy containing
nickel, chromium, cobalt and/or nickel-iron can be advantageously
used to do so. Typical high-temperature casting alloys are in
particular Inconel 738/Inconel 738 LC, Inconel 939/Inconel 939 LC,
Inconel 713/Inconel 713 LC, C1023, Mar M 002 and/or CM 274LC.
Casting methods (for example precision casting) allow the
manufacture of segment components 10 with very thin walls and in
very complex shapes.
It is thus for example advantageous when the combustion-chamber
walls 11, 12 have a mean thickness between 1 and 4 mm. The wall of
the combustion-chamber head 22 can be between 2 and 4 mm. It is for
example possible during shaping to integrally cast nozzles 15 for
air cooling. It is also possible to cast mounting flanges 23 on the
combustion-chamber head 22 in one piece. In principle, the
possibilities for shaping are not restricted to the features
illustrated.
The combustion-chamber walls 11, 12 of this embodiment are
contoured in a specific way: the inner combustion-chamber wall 11
has a bulge 13 which points downward in the representation selected
here. The bulge 13 thus points away from the burner axis 21. The
outer combustion-chamber wall 12 has an approximately identically
shaped bulge 14 upwards. This bulge 14 thus also faces away from
the burner axis 21.
The bulges 13, 14 are arranged here such that they approximately
correspond to the contour of the fuel flame 20 when the annular
combustion chamber is in operation.
These correlations are shown schematically in FIGS. 2B, C, where
FIG. 2B shows a longitudinal section through the annular space 30
and FIG. 2C shows a sectional view perpendicularly thereto. In the
sectional view of FIG. 2B, the fuel flame 20 is shown
schematically, extending from the injector 25 into the annular
space 30 over a length L.sub.B. The length of the entire annular
space is referred to as L. It is advantageous when L.sub.B=0.5-0.9
L applies for the length L.sub.B of the fuel flame 20. This means
that the fuel flame 20 extends over 50 to 90% of the axial extent
of the annular space.
The bulge 13 on the inner combustion-chamber wall 11 and the bulge
14 on the outer combustion-chamber wall 12 reach in the axial
direction approximately the distance by which the fuel flame 20
extends into the annular space.
In advantageous embodiments, the axial extent of the bulges 13, 14
is about 50 to 90% of the entire axial extent of the annular space.
Furthermore, it is advantageous when the width B.sub.B of the
bulges 13, 14 is about 30 to 60% of the width B of a segment
component 10, where the width B.sub.B of the bulge on the inside is
smaller than on the outside.
FIG. 2C shows the sectional view perpendicularly to the view of
FIG. 2B, from which it can also be discerned that the bulges 13, 14
are adapted approximately to the contour of the fuel flame.
In FIG. 20 an area A is shown in which the annular space height
H.sub.RA is increased by the bulges 13, 14, and an area B in which
the annular space height H.sub.RB is reduced.
An arc length U of the segment component 10 is thus made up of
A+2B. It is advantageous when the proportion of the area A is 50 to
80% of the arc length U and the proportion of the area B is 20 to
50% of the arc length U.
Furthermore, in FIG. 2C the usual radii of the combustion-chamber
walls are indicated, i.e. R.sub.i and R.sub.a, where it can be
discerned that bulges 13, 14 are in part outside of R.sub.a or
inside of R.sub.i. The usual (conventional) annular space height
H.sub.konv thus corresponds to R.sub.a-R.sub.i.
Advantageous embodiments have bulges 13, 14 for which applies:
H.sub.RA=1.1-1.5 H.sub.konv. This means that the height of the
combustion space in the area of the bulges 13, 14 is extended by 10
to 50% compared with the conventional design.
It is also advantageous when in the area B, i.e. in areas without
bulges 13, 14, the following applies: H.sub.RA=0.7-0.9 H.sub.konv.
This means that the height of the combustion space in the area
outside the bulges 13, 14 is 70 to 90% of the usual height.
If several of these segment components 10 are now connected to one
another, an annular combustion chamber is formed of which the
annular space height H.sub.R in the circumferential direction is
variable. Segment components 10 are for example connected to one
another by laser or electron beam welding, where the energy input
per unit length is minimized. A suitable ductile filler can be used
for welding (IN625 or Polymet 972).
An annular combustion chamber assembled in this manner is shown in
FIG. 3. For reasons of clarity, only six segment components 10 are
used here to form an annular space 30. Areas A with a greater
annular space height H.sub.RA alternate with areas B with a lower
annular space height H.sub.RB along the circumference, such that
the combustion-chamber walls 11, 12 form a kind of wavelike
structure.
The fuel flames 20 (not shown here) are in each case in the
expanded areas A. Narrowed areas B are located between the fuel
flames 20. This leads to each fuel flame 20 being able to burn
practically in its own combustion space. Perturbations in one area
of the annular space 30 cannot spread so easily inside the entire
annular space 30 because of the narrowed sections in the areas
B.
Air can also be routed in the areas B between the injectors 25 with
less heavy deflection from the compressor to the turbine 40, so
that the pressure loss on this flow path drops.
The embodiment described however also has advantageous effects
outside the annular space 30, since the turbine cooling air K too,
which is routed outside the annular space, is influenced by the
contouring of the combustion-chamber walls 11, 12.
Here the pressure loss during the passage of the turbine cooling
air K from the compressor outlet past the combustion chamber to the
inlet into the cooling system is determined in this way by the flow
guidance. If the turbine cooling air K has to be repeatedly (in
particular radially) deflected and accelerated (and then
decelerated again), then the pressure loss increases. In the burner
axis 21, only little turbine cooling air K flows past the burner
and the mixed air hole in the direction of the turbine, so the
pressure loss there is not so crucial.
Between the burners in the present embodiment, the
combustion-chamber head 22 is designed such that the turbine
cooling air K is not first heavily deflected radially outwards and
inwards. These are the areas B between the bulges 13, 14, but on
the respective outer faces of the annular space 30. Radial
deflection is followed by a deflection in the axial direction.
There is thus in area B a minor deflection into the much deeper
annuli around the combustion chamber which is narrower at this
point. The flow of turbine cooling air K is schematically shown in
FIG. 3.
With appropriate flow guidance, pressure losses are lower. The
pressure loss is reduced by the indentation between the burners.
Due to the deeper annuli, the turbine cooling air K has, in
comparison with the usual gap flow, less contact to the hot
combustion-chamber wall and is thus supplied colder to the turbine,
which improves the cooling effect inside the turbine.
In all, the total pressure loss can be reduced, lowering the fuel
consumption. In addition, less air flows between the injectors 25
into the area of the combustion-chamber head 22 than at the
position of the injectors 25, so that sufficient air is available
for transfer into the turbine 40 at these circumferential
positions.
Moreover, the bulges 13, 14 lead to a more even temperature
distribution in the circumferential direction inside the
combustion-chamber walls 11, 12, which has a positive effect on the
service life of the annular combustion chamber. In the areas A in
which the fuel flame 20 is located, the combustion-chamber wall 11,
12 is, due to the bulges 13, 14, relatively far away from the fuel
flame 20. In the areas B between the fuel flames 20, the
combustion-chamber walls 11, 12 are closer together, since the
annular space height H.sub.R is lower here. Without the bulges 13,
14, the wall areas of the combustion-chamber walls 11, 12 closest
to the fuel flame 20 would be hotter than other areas. For these
reasons, it is not necessary to use so much cooling air in the area
A. The cooling air thus saved is available for measures to reduce
the exhaust emissions.
As can be discerned in FIG. 3, the inner combustion-chamber wall 11
and the outer combustion-chamber wall 12 have a wavy structure if
they are assembled from segment components 10, for example in
accordance with FIG. 2. This wavy structure permits an easier
compensation for thermal and/or mechanical stresses in the
combustion-chamber walls 11, 12 than would be the case in annular
spaces with circular cross-sections in the circumferential
direction.
If it seems necessary (for example in larger aircraft engines), the
segment components 10 can be provided with a thermal barrier
coating.
If a ductile filler material is used, it is not necessary, in the
case of subsequent laser drilling of the annular combustion
chamber, to take account of the positions of longitudinal welds
between the segment components 10.
FIG. 4 shows a further embodiment of a segment component 10. In
principle it has the same functions and properties as the
previously described segment component 10, so that the appropriate
description can be referred to.
Unlike the substantially rectangular bulges 13, 14 in the
embodiment according to FIG. 2, the bulges 14 here are arranged in
the shape of the fuel flame 20 from the combustion-chamber head 22
in the direction of the turbine 40 (not shown here). The bulge 13
has a rather low width in the vicinity of the combustion-chamber
head 22, which steadily increases and then decreases again.
In principle, the casting method can also be used to provide other
shapes for bulges that can be adapted to a certain intended use.
The use of the aforementioned materials and the casting method in
particular make it possible to shape the bulges 13, 14
selectively.
FIGS. 2, 3 and 4 show embodiments in which two combustion-chamber
walls 11, 12 are opposite. These segment components 10 thus have a
substantially U-shaped arrangement, since the combustion-chamber
walls 11, 12 are connected by the combustion-chamber head 22 cast
in one piece with them.
It is however also possible in principle that a segment component
10 has only an outer or an inner part of the annular combustion
chamber. FIGS. 4 and 5 show an embodiment of a segment component 10
having only an outer combustion-chamber wall 12. Like the
previously described embodiments, this segment component 10 too has
a bulge 14 pointing away from the burner axis 21. To make clear the
use of this segment component 10, FIG. 4 shows in dashed lines the
fuel flame 20 and the burner axis 21.
With this embodiment too and with a corresponding segment component
10 for the inner combustion-chamber wall 11, an annular combustion
chamber can be designed as shown in FIGS. 6A, B.
To do so, at least two segment components 10' are connected, in
particular welded, to an inner full ring structure 31. Furthermore,
two segment components 10'' are connected, in particular welded, to
an outer full ring structure 32. FIG. 6A shows the two full ring
structures 31, 32 which, for reasons of simplicity, have only six
segment components 10. Then the inner full ring structure 31 and
the outer full ring structure 32 are connected to a
combustion-chamber head structure 43 as shown in FIG. 6B.
LIST OF REFERENCE NUMERALS
10 Segment component 11 Inner combustion-chamber wall 12 Outer
combustion-chamber wall 13 Bulge, inner combustion-chamber wall 14
Bulge, outer combustion-chamber wall 15 Nozzle for cooling air 20
Fuel flame 21 Burner axis 22 Combustion-chamber head 23 Mounting
flange 24 Device for arrangement of a burner 25 Injector for fuel
30 Annular space 31 Inner full ring structure 32 Outer full ring
structure 40 Intake area of turbine K Turbine cooling air H.sub.RA
Area of greater annular space height H.sub.RB Area of lower annular
space height H.sub.R Annular space height H.sub.konv Usual
(conventional) annular space height R.sub.i Radius of inner
combustion-chamber wall R.sub.a Radius of outer combustion-chamber
wall B Width of segment component B.sub.B Bulge width L.sub.B
Length of fuel flame L Length of fuel chamber U Arc length of a
segment component
* * * * *